U.S. patent number 8,393,158 [Application Number 12/000,066] was granted by the patent office on 2013-03-12 for low shock strength inlet.
This patent grant is currently assigned to Gulfstream Aerospace Corporation. The grantee listed for this patent is Timothy R. Conners. Invention is credited to Timothy R. Conners.
United States Patent |
8,393,158 |
Conners |
March 12, 2013 |
Low shock strength inlet
Abstract
A supersonic inlet with a cowl lip may be configured to capture
the conic shock and exhibit a zero or substantially zero cowl
angle. The inlet may be configured to employ a relaxed isentropic
compression surface and an internal bypass. The nacelle bypass may
prevent flow distortions, introduced by the capture of the conic
shock, from reaching the turbomachinery, thereby allowing the cowl
angle to be reduced to zero or substantially zero. Such a cowl
angle may reduce the inlet's contribution to the overall sonic boom
signature for a supersonic aircraft while allowing for an increase
in engine pressure recovery and a subsequent improvement in
generated thrust by the engine.
Inventors: |
Conners; Timothy R.
(Statesboro, GA) |
Applicant: |
Name |
City |
State |
Country |
Type |
Conners; Timothy R. |
Statesboro |
GA |
US |
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Assignee: |
Gulfstream Aerospace
Corporation (Savannah, GA)
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Family
ID: |
40581291 |
Appl.
No.: |
12/000,066 |
Filed: |
December 7, 2007 |
Prior Publication Data
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Document
Identifier |
Publication Date |
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US 20090107557 A1 |
Apr 30, 2009 |
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Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
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60960986 |
Oct 24, 2007 |
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Current U.S.
Class: |
60/767;
137/15.1 |
Current CPC
Class: |
F02C
7/042 (20130101); F02K 7/16 (20130101); F02C
7/045 (20130101); B64D 33/06 (20130101); F02K
3/025 (20130101); B64D 29/00 (20130101); B64D
33/02 (20130101); Y10T 137/0536 (20150401); F05D
2220/80 (20130101); B64D 2033/0273 (20130101); B64D
2033/026 (20130101); Y02T 50/60 (20130101); Y02T
50/672 (20130101); Y10T 137/0368 (20150401) |
Current International
Class: |
F02K
7/08 (20060101) |
Field of
Search: |
;60/767,768,226.1
;137/15.1,15.2 ;244/53B |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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1527997 |
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May 2005 |
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EP |
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879956 |
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Oct 1961 |
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GB |
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885908 |
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Jan 1962 |
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GB |
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2 070 139 |
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Sep 1981 |
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GB |
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2149456 |
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Jun 1985 |
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GB |
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09-291850 |
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Nov 1997 |
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JP |
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WO 2008/045108 |
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Apr 2008 |
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WO |
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WO 2009/055041 |
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Apr 2009 |
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WO |
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WO 2009/085380 |
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Jul 2009 |
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WO |
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Other References
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University; Jul. 1, 1974-Sep. 30, 1974, pp. 1-22. cited by
applicant .
"Conceptual Design, Integration, and Development Plan for an
Efficient Low Sonic Boom Propulsion System Employing Advanced
Supersonic Engine Cycles;" Integrated Inlet-Propulsion Systems
Study Final Report; Conners et al.; Aug. 27, 2007; pp. 1-84. cited
by applicant .
"Wind Tunnel Testing of an Axisymmetric Isentropic Relaxed External
Compression Inlet at Mach 1.97 Design Speed;" American Institute of
Aeronautics and Astronautics, Inc.; Conners et al.; Jul. 8-11,
2007; pp. 1-12. cited by applicant .
"Dynamic Analysis of Wind Tunnel Data from an Isentropic Relaxed
Compression Inlet;" American Institute of Aeronautics and
Astronautics; Tacina et al.; Jul. 8-11, 2007; pp. 1-22. cited by
applicant .
"Supersonic Inlet Shaping for Dramatic Reductions in Drag and Sonic
Boom Strength;" American Institute of Aeronautics and Astronautics,
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.
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Airplane Group, HSCT Designs for Reduced Sonic Boom. cited by
applicant .
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Minimization Revisited. cited by applicant .
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Communications on Pure and Applied Math, vol. V, 301-348). cited by
applicant .
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applicant .
AIAA 68-159, 1968, A. George, Reduction of Sonic Boom by Azimuthal
Redistribution of Overpressure. cited by applicant .
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Attainable Sonic-Boom Overpressure and Design Methods of
Approaching This Limit. cited by applicant .
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Sonic-Boom Theory With Wind-Tunnel and Flight Measurements. cited
by applicant .
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Effects on the Sonic Boom of Large Airplanes, Jun. 1965. cited by
applicant .
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Class of Bodies that Generate Far-Field Sonic-Boom Shock Strength
and Impulse Independent of Body Length and Volume. cited by
applicant .
Gokhale, et al., "Numerical computations of supersonic inlet flow,"
International Journal for Numerical Methods in Fluids 2001; 36:
597-617. cited by applicant .
Emami, et al., "Experimental Investigation of Inlet-Combustor
Isolators for a Dual-Mode Scramjet at a Mach Number of 4," NASA
Technical Paper 3502, May 1995. cited by applicant .
NASA SP-255, 1971, Edited by I. Schwartz, Third Conference on Sonic
Boom Research. cited by applicant .
NASA Technical Note TN D-7218, 1973, H. Carlson, Application of
Sonic-Boom Minimization Concepts in Supersonic Transport Design.
cited by applicant .
NASA Technical Note TN D-7842, 1975, C. Darden, Minimization of
Sonic-Boom Parameters in Real and Isothermal Atmospheres. cited by
applicant .
NASA Technical Paper 1348, 1979, C. Darden, Sonic-Boom Minimization
With Nose-Bluntness Relaxation. cited by applicant .
NASA Technical Paper 1421, 1979, R. Mack, Wind-Tunnel Investigation
of the Validity of Sonic-Boom-Minimization Concept. cited by
applicant .
NASA Technical Note TN D-7160, 1973, L. Hunton, Some Effects of
Wing Planform on Sonic Boom. cited by applicant .
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Sonic Boom Pressure Signatures by the Waveform Parameter Method.
cited by applicant .
Paper abstract, K. Plotkin, Wyle Laboratories, Sonic Boom
Minimization: Myth or Reality, Jan. 24, 1975. cited by applicant
.
NASA SP-147, 1967, A.R. Seebass, Sonic Boom Research. cited by
applicant .
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Sonic Boom Research. cited by applicant .
U.S. Office Action issued Mar. 22, 2012 in U.S. Appl. No.
12/257,982. cited by applicant .
AIAA 2006-30, 2006, Conners et al., Supersonic Inlet Shaping for
Dramatic Reductions in Drag and Sonic Boom Strength; pp. 1-24.
cited by applicant .
European Extended Search Report dated Sep. 21, 2012; Application
No. 08842021.1. cited by applicant .
European Extended Search Report dated Sep. 21, 2012; Application
No. 08867105.4. cited by applicant .
USPTO "Non-Final Office Action" mailed Oct. 5, 2012; U.S. Appl. No.
12/257,982, filed Oct. 24, 2008. cited by applicant.
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Primary Examiner: Wongwian; Phutthiwat
Attorney, Agent or Firm: Ingrassia Fisher & Lorenz,
P.C.
Parent Case Text
This application claims priority to co-pending U.S. Provisional
Patent Application 60/960,986, filed Oct. 24, 2007, and entitled
"Supersonic Nacelle," which is hereby incorporated by reference in
its entirety.
Claims
What is claimed is:
1. A supersonic inlet: comprising: a center body having a leading
edge that generates a first shock wave when positioned in
supersonic air flow, the center body comprising a compression
surface positioned downstream from the leading edge of the center
body; a cowl lip spatially separated from the center body such that
the cowl lip and the compression surface of the center body define
an external compression inlet opening that receives supersonic air
flow, at least a portion of the compression surface extending
upstream of the cowl lip, and the cowl lip being separated from the
compression surface by a second distance; and a bypass splitter
disposed between the cowl lip and the center body to form a bypass
air path that bypasses a bypass flow portion to an engine, wherein
the compression surface generates a second shock wave that, during
operation of the supersonic inlet at a predetermined cruise speed,
extends from the compression surface intersects the first shock
wave at a first point that is spatially separated from the
compression surface by a first distance, the first distance being
less than the second distance such that the inlet captures the
first shock wave.
2. The supersonic inlet of claim 1, wherein the compression surface
includes at least one curved section configured to generate
isentropic compression.
3. The supersonic inlet of claim 2, wherein the isentropic
compression generated by the curved section is characterized by a
series of Mach lines where, during operation of the supersonic
inlet at the predetermined cruise speed, at least a plurality of
the Mach lines do not focus on the first point.
4. The supersonic inlet of claim 1, wherein the bypass divides an
incoming airflow the inlet into: a primary flow portion received by
the engine; and the bypass flow portion, the bypass flow portion
comprising at least a portion of flow distortion created by the
intersection of the first and second shock waves at the first
point.
5. The supersonic inlet of claim 1, wherein the cowl lip is not
aligned with a flow angle adjacent to the cowl lip.
6. The supersonic inlet of claim 5, further comprising an inlet
axis aligned with a longitudinal axis of the center body, wherein
the cowl lip is substantially parallel with the inlet axis.
7. The supersonic inlet of claim 1, wherein an opening of the inlet
is an axisymmetric inlet opening.
8. The supersonic inlet of claim 1, wherein an opening of the inlet
is a non-axisymmetric inlet opening.
9. The supersonic inlet of claim 1, wherein the cowl lip has a cowl
angle that is substantially zero.
10. The supersonic inlet of claim 1, wherein the second shock wave
is generated substantially near an entrance of the external
compression inlet opening.
11. A supersonic propulsion system: comprising: an engine having an
air intake and an exhaust system; a subsonic diffuser section
coupled to the air intake of the engine; and a supersonic
compression section coupled to the subsonic diffuser and including
a compression surface, a bypass splitter that bypasses a bypass
flow portion to the engine, and a cowl lip, the cowl lip being
spatially separated from the compression surface such that the cowl
lip and the compression surface define an external compression
inlet opening that receives supersonic air flow, at least a portion
of the compression surface extending upstream of the cowl lip;
wherein, when positioned in the supersonic air flow, the
compression surface generates a first shock wave off a leading edge
of the compression surface and a second shock wave, the second
shock wave extending from the compression and intersecting the
first shock wave at a first point that is spatially located at a
distance that is less than a distance between the compression
surface and the cowl lip such that the external compression inlet
opening captures the first shock wave.
12. The supersonic propulsion system of claim 11, wherein the
bypass splitter divides an airflow entering the inlet opening into
a primary flow portion received by the engine and the bypass flow
portion, the bypass flow portion comprising at least a portion of
flow distortion created by the intersection of the first and second
shock waves at the first point.
13. The supersonic propulsion system of claim 12, wherein the
subsonic diffuser and the bypass splitter are configured to diffuse
the primary flow portion to a subsonic condition suitable for the
engine.
14. The supersonic propulsion system of claim 11, wherein the cowl
lip is not aligned with a flow angle adjacent to the cowl lip.
15. The supersonic propulsion system of claim 14, further
comprising an inlet axis, wherein the cowl lip is substantially
aligned with the inlet axis.
16. The supersonic propulsion system of claim 11, wherein an
isentropic compression generated by the compression section is
characterized by a series of Mach lines where, during operation of
the system at a predetermined cruise speed, at least a plurality of
the Mach lines do not focus on the first point.
17. The supersonic propulsion system of claim 11, wherein the cowl
lip has a cowl angle that is substantially zero.
18. The supersonic propulsion system of claim 11, wherein the
second shock wave is generated substantially near an entrance of
the external compression inlet opening.
Description
FIELD OF INVENTION
The embodiments of the present invention relate generally to
supersonic inlets for supersonic aircraft and more particularly to
supersonic inlets configured to reduce the supersonic inlet's
contribution to the aircraft's sonic boom signature.
BACKGROUND OF THE INVENTION
Gas turbine engines can propel aircraft at supersonic speeds.
However, the gas turbine engines generally operate on subsonic air
flow in the range of about Mach 0.3 to 0.6 at the upstream face of
an engine. The inlet of the engine functions to decelerate the
incoming airflow to a speed compatible with the requirements of the
gas turbine engine. In order to do this, the inlet has a
compression surface and a corresponding flow path, used to
decelerate the supersonic flow into a strong terminal shock. A
diffuser further decelerates the resulting flow from the strong
terminal shock to a speed corresponding to the requirements of the
gas turbine engine.
A measurement of inlet operation efficiency is the total pressure
lost in the air stream between the entrance side and the discharge
side of the inlet. The total pressure recovery of an inlet is
defined by a ratio of the total pressure at the discharge to the
total pressure at the free stream. Maximizing the total pressure
recovery leads to maximizing gross engine thrust, thus improving
the performance of the propulsion system. Traditional inlet design
methods have aimed at maximizing total pressure recovery. This
traditional approach, however, often results in a complex inlet
design with high drag.
A traditional approach to supersonic inlet design typically employs
shock-on-lip focusing. As understood by those of skill in the art,
shock-on-lip focusing involves designing a compression surface
configuration of an external compression inlet such that the
inlet-generated shocks (that occur at a supersonic design cruise
speed) meet at a location immediately forward of the cowl highlight
or the cowl lip. The advantages of shock-on-lip focusing include
better pressure recovery and low flow spillage drag.
Also, when using shock-on-lip focusing, the cowl lip angle of the
cowling may be aligned with the local supersonic flow in the
vicinity of the terminal shock in order to prevent formation of an
adverse subsonic diffuser flow area profile or a complex internal
shock structure in the lip region. If this is not done, a complex
internal shock structure and an adverse subsonic diffuser flow area
profile may result, possibly reducing the inlet pressure recovery
and flow pumping efficiency, as well as undermining diffuser flow
stability.
As understood in the art, as supersonic design speed increases, so
will the amount of compression necessary to decelerate the flow to
a fixed terminal shock Mach number. Additional compression requires
more flow-turning off of the inlet axis, resulting in a
corresponding increase in the cowl lip angle (in order to align the
cowl lip angle with the local flow at the terminal shock). FIG. 1
schematically illustrates a side view of a conventional inlet 1.
Inlet 1 has a compression surface 10 and a cowl lip 11. Cowl lip 11
is positioned such that both an initial shock and a terminal shock
from compression surface 10 meet at a point before the cowl lip 11.
A cowl lip angle 12 is formed when the cowl lip 11 is aligned with
the local flow. As mentioned, when the supersonic design speed
increases, the amount of compression needed to decelerate the flow
to a fixed terminal shock Mach number also increases, resulting in
an increase in cowl lip angle. Any increase in cowl lip angle
results in additional inlet frontal area, which increases inlet
drag as speed increases. This adverse trend is a key reason why
conventional external compression inlets lose viability at high
supersonic Mach numbers.
One way to control lip drag, as discussed in U.S. Pat. No.
6,793,175 to Sanders, involves configuring the inlet to minimize
the shape and size of the cowl. The configuration of the inlet
initially resembles a circumferential sector of an axisymmetric
intake, but switches the location of compression surface to the
outer radius and disposes the cowling on the inner radius in a
higher performance, 3-D geometry. The fact that the cowl is located
on the inner radius reduces the physical arc of the cowl. Problems
with this method include the aircraft integration challenges
created by the 3-D geometry, such as the fact that the
cross-sectional shape may be more difficult to integrate from a
packaging perspective compared to an equivalent axisymmetric design
for podded propulsion systems. In addition, the complex inlet shape
is likely to create complex distortion patterns that require either
large scale mitigating techniques in the subsonic diffuser or the
use of engines with more robust operability characteristics.
Another way to control drag by reducing the cowl lip angle is based
on decreasing the flow turn angle by increasing the inlet terminal
shock Mach number. The improvement in drag reduction is often
negated by the reduction in pressure recovery resulting from the
stronger terminal shock. In addition, increasing the terminal shock
Mach number at the base of the shock also encounters significant
limitations in practice due to viscous flow effects. Higher
terminal shock Mach numbers at the base of the shock aggravate the
shock-boundary layer interaction and reduce shock base boundary
layer health. The increase in shock strength in the base region
also reduces inlet buzz margin, reducing subcritical flow
throttling capability. Additionally, the increase in terminal shock
Mach number will most likely require a complex boundary layer
management or inlet control system.
Inlet compression surfaces are typically grouped into two types:
straight or isentropic. A straight surface has a flat ramp or conic
sections that produce discrete oblique or conic shocks, while an
isentropic surface has a continuously curved surface that produces
a continuum of infinitesimally weak shocklets during the
compression process. Theoretically, a traditional isentropic
compression surface can have better pressure recovery than a
straight surface designed to the same operating conditions, but
real viscous effects can reduce the overall performance of the
isentropic surface inlets and result in poorer boundary layer
health.
SUMMARY OF THE INVENTION
In accordance with one embodiment of the invention, a supersonic
inlet may include a leading edge configured to generate a first
shock wave, a compression surface positioned downstream of the
leading edge, and a cowl lip spatially separated from the
compression surface such that the cowl lip and the compression
surface define an inlet opening for receiving a supersonic flow.
The supersonic inlet may also include a bypass splitter disposed
between the cowl lip and the center body to form a bypass. The
compression surface may also be configured to generate a second
shock wave, which during operation of the supersonic inlet at a
predetermined cruise speed, extends from the compression surface to
intersect the first shock wave at a first point spatially separated
from the compression surface by a distance less than the distance
separating the compression surface and the cowl lip such that the
inlet captures the first shock wave.
In another embodiment of the invention, a supersonic propulsion
system may be configured to include an engine having an air intake
and an exhaust system, a subsonic diffuser section coupled to the
air intake of the engine, and a supersonic compression section
coupled to the subsonic diffuser and including a compression
surface, a bypass splitter, and a cowl lip. The cowl lip may be
spatially separated from the compression surface such that the cowl
lip and the compression surface define an inlet opening for
receiving a supersonic flow. The compression surface may also be
configured to generate a first shock wave off a leading edge of the
compression surface and a second shock wave such that the second
shock wave extends from the compression surface to intersect the
first shock wave at a first point located between the compression
surface and the cowl lip, such that the inlet opening captures the
first shock wave.
Another example of an embodiment of the invention may include the
method of decelerating a supersonic flow for a supersonic
propulsion system where the method includes cruising at a
predetermined supersonic speed, receiving a supersonic flow in an
inlet opening of an supersonic inlet of the supersonic propulsion
system, generating a first shock wave, generating a second shock
wave that intersects the first shock wave, receiving the first
shock wave, during operation of the inlet at a predetermined
supersonic speed, in the inlet opening, and splitting a subsonic
flow into a primary flow portion and a bypass flow portion, whereby
the bypass flow portion separates a substantially all flow
distortion introduced when the inlet opening receives the first
shock wave.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 schematically illustrates a side view of a conventional type
of inlet for a supersonic aircraft.
FIG. 2 schematically illustrates a side elevation view of a
supersonic aircraft inlet entrance.
FIG. 3 schematically illustrates a side view of an inlet in
accordance with an embodiment of the invention.
FIG. 4 illustrates a Mach color Computational Fluid Dynamics (CFD)
solution of an inlet with a conventional cowl.
FIG. 5 illustrates a Mach color CFD solution of an external
compression inlet with a zero-angle cowl in accordance with an
embodiment of the invention.
DETAILED DESCRIPTION OF THE INVENTION
The present disclosure will now be described more fully with
reference to the Figures in which various embodiments of the
invention are shown. The subject matter of this disclosure may,
however, be embodied in many different forms and should not be
construed as being limited to the embodiments set forth herein.
An embodiment of the invention may include a supersonic inlet for
supersonic aircraft that is configured to reduce the inlet's
contribution to a supersonic aircraft's sonic boom signature. To
accomplish this, embodiments of the invention may position the cowl
lip of the inlet such that the inlet captures the initial conic
and/or oblique shock within the intake plane, preventing the conic
shock energy or discontinuity from merging with the shocks
generated by the airframe during supersonic flight. It is also
contemplated that the cowl angle of the nacelle may be reduced to
zero or substantially zero in order to reduce the contribution of
cowl shock and cowl drag on the overall signature of a supersonic
aircraft.
When designing an inlet in accordance with an embodiment of the
invention, a relaxed isentropic compression surface may be used. As
discussed in commonly owned U.S. patent application Ser. No.
11/639,339, filed Dec. 15, 2006 (entitled "Isentropic Compression
Inlet for Supersonic Aircraft"), which is hereby incorporated by
reference in its entirety, a reduction in cowl angle may be
achieved by designing an inlet to employ a relaxed isentropic
compression surface such that the cowl angle may be reduced. A
"relaxed isentropic compression" surface is an isentropic
compression surface where a plurality of Mach lines do not focus on
the focus point where the initial shock and the terminal shock
meet. This lack of Mach line focusing may be configured to produce
a total level of compression less than the level of compression
generated by a conventional isentropic compression surface designed
to the same criteria. The relaxed isentropic compression surface
may be configured to increase terminal shock Mach number in the
region of the cowl lip (creating the mechanism that reduces flow
angle at the lip), but retains a reasonable terminal shock Mach
number along the remainder of the shock, including the base region
of the terminal shock (preserving a reasonable overall pressure
recovery characteristic and good shock stability). Such an
arrangement may significantly reduce the local flow angle at the
cowl lip, leading to a reduction in cowling angle and a substantial
improvement in performance and a reduction in shock strength.
FIG. 2 schematically illustrates a side view cross section of a
relaxed isentropic external compression inlet 100 configured using
shock-on-lip focusing as disclosed in U.S. patent application Ser.
No. 11/639,339. The inlet 100 includes a compression surface 110
with an initial straight surface 140 at an initial turn angle 110a.
The compression surface 110 includes a second compression surface
111 comprising a curved section 112 and a straight section 113. The
compression surface 110 transitions into a shoulder 130, which
defines the throat 135, the narrowest portion of the inlet 100 flow
path. The inlet 100 also includes cowl lip 120 positioned at a cowl
angle 110b measured off the centerline of the inlet 100. Although
only the curved section 112 of the second compression surface 111
generates isentropic compression, the entire compression surface
110 is referred to herein as a relaxed isentropic compression
surface. For comparison, an example of a traditional isentropic
compression surface 160 is shown in a dashed line. After the flow
reaches throat 135, subsonic diffuser 150 provides a divergent flow
path delivering subsonic flow to the engine.
The inlet 100 first generates an initial shock 200 as the air flow
in region B travels in direction A and encounters the compression
surface 110 of inlet 100. The compression surface 10 may be
configured to generate a terminal shock 210, having a base 210a
adjacent to the compression surface 110. As shown in FIG. 2, the
initial shock 200 and the terminal shock 210 are focused at a shock
focus point 230. A cowl shock 220 is shown extending upward off the
cowl lip 120. The relaxed isentropic compression surface allows for
significant tailoring of the terminal shock 210 such that the outer
radial region of the shock is nearly orthogonal to the inlet
centerline. By shaping the terminal shock using relaxed
compression, the cowl lip 120 may be aligned with the local flow
angle in this outer radial region of the shock, greatly reducing
the cowl lip angle. In addition, discrete adverse flow features,
such as secondary shock formation or flow separation, may be
reduced at the cowl lip region.
Although the cowl angle may be greatly reduced when using a relaxed
isentropic compression inlet in accordance with FIG. 2, the cowl
lip is still aligned with the local flow angle in the outer radial
region of the terminal shock directly in front of the cowl lip. As
would be understood by those of skill in the art, reducing the cowl
angle 110b, from the angle shown in FIG. 2 to zero or substantially
zero may result in flow distortion in the diffuser which may
increase when the cowling angle no longer aligns with the local
flow in the vicinity of the terminal shock. This condition may
generate secondary shocks and adverse pressure fields in the
vicinity of the cowl lip, which can introduce strong tip radial
blockage defects in the flow seen by the engine at the fan face.
Further, simply reducing the cowl angle 110b to zero or
substantially zero may also create temporal flow instability within
the diffuser, potentially resulting from the flow disturbances
created in the outer radial region which may initiate and sustain
diffuser flow resonance. Such resonance may adversely affect
performance and potentially damage the inlet and the engine.
Additionally, a simple reduction in cowl angle may be ineffective
in controlling aft cowling drag, or drag on the nacelle aft of the
cowl lip resulting from any increase in nacelle diameter as the
nacelle profile encompasses the engine. This increase in nacelle
diameter may cause a sharper gradient in the surface angle of the
cowling as the maximum nacelle diameter is approached.
Furthermore, when the cowl lip is positioned to capture the initial
or conic shock and the terminal shock in accordance with
embodiments of the invention, flow instabilities internal to the
inlet may be introduced. As understood by those of skill in the
art, the capture of the conic and terminal shocks may decrease the
predictability of the post terminal shock flow environment and
introduce flow separation on the inside cowl surface and produce
unwanted flow dynamics.
Embodiments of the invention may be configured to mitigate the
above-discussed adverse effects of zero cowl angle and conic and
terminal shock capture by employing a flow bypass system to
separate and isolate the outer radial flow captured by an inlet and
bypass that separated flow around the engine. Embodiments of the
invention may use the nacelle bypass design as described in
commonly owned U.S. Patent Application No. 60/960,986, filed Oct.
24, 2007 (entitled "Supersonic Nacelle"), which is hereby
incorporated by reference in its entirety.
By combining initial shock capture, an internal bypass, and a zero
cowl angle, embodiments of the invention may be configured to
reduce spillage-related drag and cowl shock strength by capturing
the strength of the initial conic shock and the terminal shock
internal to the inlet. More specifically, capture of the conic and
terminal shocks may permit the shock energy or discontinuity to be
retained within the nacelle flow paths, preventing the shock from
merging with shocks generated by the airframe during supersonic
flight and contributing to the overall sonic boom signature. The
use of a nacelle bypass flow path may be configured to provide a
separation, isolation, and disposal mechanism for the resulting
spatial and temporal flow defects that may be produced by shock
capture and zero cowl angle, leaving a primary flow path available
for use by the engine.
FIG. 3 schematically illustrates a cross-sectional view of an inlet
300 in accordance with an embodiment of the invention. Supersonic
inlet 300 includes a center body 310 with a relaxed isentropic
compression surface 320 and a leading edge 325. It should be
understood that, while a relaxed isentropic compression surface is
shown and described with reference to FIG. 3, other compression
surfaces, such as a fully isentropic surface or a straight surface
compression surface, may be used. Inlet 300 also includes a cowl
lip 330 and a bypass splitter 340 in order to form a nacelle bypass
350. A bypass strut 360 and a primary strut 370 (which are only
shown on the bottom of the nacelle 300 and have been removed from
the top of the nacelle 300 for clarity) may provide structural
support to the inlet, producing a stiff, strong, and lightweight
nacelle structure, while maximizing the internal nacelle volume. As
discussed in U.S. Patent Application No. 60/960,986, the bypass
strut 360 can also be used to tailor the direction and the amount
of airflow depending on local blockage characteristics within the
bypass region.
As shown in FIG. 3, the inlet structure and arrangement may be
configured such that the cowl lip angle is extremely small or even
reduced to zero. As would be understood by those of skill in the
art, a zero or substantially zero cowl lip angle reduces the
strength of the cowl shock due to reductions in the projected
surface area exposed to the freestream flow. Although the thickness
of the cowl lip may include some finite amount of material required
to build the cowl lip, the cowl lip structure may be extremely
thin, depending on materials and application. It is contemplated
that the nacelle wall thickness may grow inward moving aft along
the internal flowpath, providing the volume necessary to
incorporate structure while maintaining the uniform external
diameter surface shape.
By employing a zero or substantially zero cowl lip angle, with
reference to a inlet axis 365, the region C may grow, especially if
the nacelle is configured to fully encompass the engine without
significant growth or contraction in the outer diameter of the
nacelle. Such a configuration may reduce or eliminate the typical
sharp growth of the outer diameter of the nacelle aft of the cowl
lip as the nacelle encompasses the engine. As understood by those
of skill in the art, a more cylindrical shape of uniform outer
diameter may significantly reduce cowling drag and cowl shock
strength.
In accordance with embodiments of the invention, the nacelle bypass
350 may be configured to handle the additional airflow that may
enter the inlet due to the larger region C. By employing the bypass
350, the inlet 300 may be configured to dispose of the excess flow,
which would alternatively spill around the exterior of the cowl
lip, creating higher drag and defeating the objective of a lower
sonic boom signature. The nacelle bypass 350 avoids these
spillage-related issues by routing the additional flow through the
nacelle and around the engine, eventually exhausting back to the
free stream.
The nacelle bypass 350 may also serve to separate the flow
distortion captured by the inlet 300. As discussed in U.S. patent
application Ser. No. 11/639,339, the use of a relaxed isentropic
compression surface 320 may generate an initial shock 400 and a
terminal shock 410, which may be focused at a point. The relaxed
isentropic compression surface may also be configured to tailor the
terminal shock 410 such that a region 420 of relaxed compression is
produced. As a result, the strong velocity gradient in the outer
radial region may generate the region 420 of flow distortion. In
accordance with embodiments of the invention, the bypass 350 may be
structured and arranged to separate the worst of the flow
distortion internal to the inlet 300 as shown as region 430. This
region 430 may include flow distortions introduced by the
intersection of the initial shock 400 and the terminal shock 410.
In addition, the region 430 may include flow distortion created by
the sharp cowl lip 330, which may produce unfavorable flow
distortion in the presence of cross-flow; for example, when the
vehicle experiences significant sideslip or angle-of-attack, or
when the vehicle is subjected to high crosswinds while operating on
the ground.
More specifically, the bypass 350 operates to split the distorted
flow in the region 430 into the bypass 350, forming a bypass flow
450, which is separated from the primary flow 440 by the splitter
340. The splitter 340 prevents the bypass flow 450 and its inherent
flow distortions from reaching the sensitive turbomachinery. The
resulting primary flow 440 may then exhibit more uniform flow that
may provide significant benefits to engine life and engine
maintenance factors and improved fan and compressor stability
margins. The primary flow 440 profile may also benefit the engine
performance by providing an increase in pressure recovery that
results from the removal of the more distorted, lower pressure flow
found in the region 430. The subsonic diffuser 380 may be
configured to further slow the primary flow 440 into a subsonic
flow suitable for use by the engine. Also, the blunt leading edge
345 of bypass splitter 340 may be configured to couple favorably
with cowl lip 330 to produce a reduced flow distortion profile for
the engine, similar to a traditional subsonic inlet.
The nacelle bypass 350 may also provide for the disposition of
residual discrete flow defects or temporal flow instabilities, such
as blockage profiles resulting from flow separation or secondary
shocks within the cowl lip area. The bypass 350 may work to
eliminate resonance coupling between tip radial and centerbody
boundary layer-related flow features that can otherwise create
adverse and strong instabilities, such as inlet buzz and other
resonance types.
In accordance with embodiments of the invention, the inlet 300 may
capture the initial conic or oblique shock 400 within the intake
plane of inlet 300. Capturing the conic shock 400 may be
accomplished by either a forward extension or movement of the
cowling or by sizing the inlet to a Mach number slightly lower than
the design point. Although capturing the conic shock 400 would
typically introduce large-scale flow instabilities from the
interaction between the conic shock and the boundary layer
immediately aft of the cowl lip, the bypass 350 may be configured
such that the conic shock 400 may be captured without significant
impact on the primary flow 440. As a result, the nacelle bypass 350
provides for a separation, isolation, and disposal mechanism for
the resulting spatial and temporal flow defects produced by conic
shock capture, leaving the primary flow path 440 significantly
unaffected.
FIG. 4 illustrates a Mach color computational fluid dynamics (CFD)
solution for an inlet 500 employing a relaxed isentropic
compression design and shock-on-lip focusing with a cowl lip placed
such that the conic shock is not captured by the inlet. FIG. 5
illustrates a Mach color computational fluid dynamics (CFD)
solution for an inlet 600 in accordance with an embodiment of the
invention. As with inlet 500, the inlet 600 employs a relaxed
isentropic compression design. However, inlet 600 includes a zero
cowl angle and is configured to capture the conic shock internal to
the inlet. FIGS. 4 and 5 represent inlets sized for a turbofan-type
engine featuring approximately 15,000 lbf of maximum takeoff thrust
and a moderate fan-to-compressor flow ratio of 3. Those areas of
the flow field disturbed by less than 0.01 Mach number unit from
the freestream Mach number value are rendered white in both FIGS. 4
and 5.
In comparison, the inlet 600 in FIG. 5 exhibits a greatly reduced
shock disturbance region 610 due to the zero-angle cowl and conic
shock capture. This may be easily seen by comparing the shock
disturbance region 510 in FIG. 4 and the shock disturbance region
610 in FIG. 5. In FIG. 4, a large region 510 of disturbance is
shown extending out and away from much of the forward nacelle
surface. This indicates that the cowl shock 520, in FIG. 4, is much
stronger that the cowl shock 620, in FIG. 5. The strong cowl shock
520 will propagate away from the nacelle and eventually merge with
shocks generated by aircraft airframe. In FIG. 5, however, a
relatively thin cowl shock disturbance 610 extends out and away
from only the very tip of the nacelle adjacent to the zero-angle
cowl lip. This is indicative of a much weaker cowl shock 620 that
will contribute little to the overall sonic boom signature.
Also illustrated in FIGS. 4 and 5, the reduction in spillage may be
seen for inlet 600 over inlet 500. As would be appreciated by one
of skill in the art, the flow spillage 630 shown in FIG. 5 for the
inlet 600 is significantly less than the amount of flow spillage
530 shown in FIG. 4 for the inlet 500. Specifically, FIG. 5 shows
minimal spillage close to the cowl lip, indicated by a
significantly reduced cowl shock strength. For inlet 600, these
reductions in shock strength directly reduce the inlet's
contribution to a sonic boom signature for a supersonic aircraft
employing inlet 600. As one of ordinary skill in the art will
appreciate, the capture of the conic shock functions to virtually
eliminate the flow spillage 630 and its related contribution to
shock strength. Moreover, the lack of any significant cowling
profile (due to zero cowl angle) virtually eliminates cowl shock
and cowl drag. The reduction in flow spillage 630 also reduces
drag.
FIG. 5 also illustrates the flow distortion that is separated and
isolated from the engine face. As discussed above, the zero or
substantially zero cowl angle and the capture of the conic shock
635 and terminal shock 640 may introduce flow distortions located
in the outer radial region of the inlet. Although the bypass
splitter 340 (shown in FIG. 3) is not shown in FIG. 5, the flow
distortion 650 adjacent to the cowl lip and the outer surface of
the diffuser walls illustrates adverse flow characteristics that
could be detrimental to the operability, performance, and life of
the fan blades at an engine face. As discussed above, these adverse
flow characteristics may be separated and isolated by the bypass
340.
It is contemplated that the invention could be applied to other
air-breathing propulsion systems configured for supersonic flight.
These propulsion systems could employ conventional turbojet and
turbofan engines, combined cycle engines, ramjets, or scramjets.
Propulsion systems employing variable cycle engine features, such
as fladed turbomachinery, may also be used. In addition, inlets
designed according to the disclosed technology may be axisymmetric,
two-dimensional, or three-dimensional in their intake and diffuser
design. It is also contemplated that embodiments of the invention
may be applied to other types of compression inlets, such as a
mixed compression inlet.
The foregoing descriptions of specific embodiments of the invention
are presented for purposes of illustration and description. They
are not intended to be exhaustive or to limit the invention to the
precise forms disclosed. Obviously, many modifications and
variations are possible in view of the above teachings. While the
embodiments were chosen and described in order to best explain the
principles of the invention and its practical applications, thereby
enabling others skilled in the art to best utilize the invention,
various embodiments with various modifications as are suited to the
particular use are also possible. The scope of the invention is to
be defined only by the claims appended hereto, and by their
equivalents.
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